Exemplary embodiments of the invention relate to a composite structure for an aircraft component, in order to form a structural element of an aircraft, as well as a method for producing the composite structure.
The invention deals, in particular, with the problem of ice forming on an aircraft and, hence, also relates to the field of ice protection systems that include both systems for preventing the formation of ice on an aircraft as well as systems for de-icing an aircraft.
Ice forms on the leading edges of wings, tails and horizontal stabilizers, when an aircraft, such as an airplane or a helicopter, flies through a cloud that contains supercooled water droplets. The ice forms, as the supercooled water droplets of clouds touch the aircraft. This contact introduces energy into the droplet and causes them to change from the liquid state to the solid state and, thus, to become ice. If a layer of ice grows, it has an adverse effect on the air flow over the surface concerned. If the layer is large enough, it can cause carrying or lifting problems or handling problems for the aircraft. In the worst case scenario the net result of such problems can be a stall and loss of lift force.
The term de-icing is defined as a process of removing the accumulation of frozen material, like snow or ice, from a surface.
The term anti-icing, i.e. the prevention of ice formation, is defined as the process of providing protection against the formation of accumulations of frozen material, like snow or ice, on the surface.
Typically de-icing operations and methods for preventing the formation of ice are carried out on the ground and include mechanical methods and de-icing by means of infrared radiation, the application of dry or liquid chemicals, salt, alcohols, or heated glycol compounds or glycol solutions by spraying.
For example, a protective layer of anti-icing fluid is applied for limited protection. For this purpose, airplanes or any other aircraft are de-iced, in particular, before take-off, especially under cold weather conditions. The de-icing of the aircraft wings both on the ground (before the flight) as well as during the flight is extremely relevant to the safety, because otherwise the formation of ice may cause the airplane to stall during the flight, so that it is more likely to crash.
Ice protection systems for preventing the formation of ice during the flight include directing hot bleed air from the engines through inner lines along the edge of the wing or applying additional heating elements that are embedded in rubber layers and mounted externally on the leading edges of aircraft parts, such as wings or propellers and helicopter rotor blades.
There are also electromechanical ice protection systems, including the so called “weeping wing” systems, where a glycol based anti-icing fluid is pumped through small openings in the wing profile. Other electromechanical ice protection systems include the so called electromechanical expulsion de-icing systems (EMEDS) systems that use a mechanical force to knock the ice off the wing surface. In this case actuators are typically installed under the skin of the structure. The actuator is moved to induce a shock wave in the protected surface. There are also hybrid systems that combine such an ice expulsion system with electric heating elements.
Known ice protection systems are described briefly below.
Pneumatic de-icing profiles, which remove ice by means of surface movements induced by air pressure, can only be used for a low speed aircraft.
Widely used are the so called “bleed air” systems, where hot air can be diverted from the engines and used for de-icing or to prevent ice formation. However, the engine bleed temperature is usually too high for aircraft components made of composite materials, such as carbon fiber-reinforced plastic (CFRP) structures. Diverting the hot air results in a degradation of the efficiency of the engine.
The heating elements that are embedded in rubber layers have drawbacks in terms of low electrical efficiency, in terms of a high weight and with regard to installing the heating elements. The heating elements can be dislodged from the aircraft during the flight. Such heating elements cannot be installed on just any surface structure. That is, the design freedom with respect to surfaces that are to be protected is extremely limited. The heating surfaces have a short service life. Such ice protection systems require high production and maintenance costs and do not lend themselves to small aircraft components, such as dynamic pressure sensors or the like.
The weeping wing systems constantly consume fluids during the flight, need nozzles on the leading edge of the wing, and pollute the environment. They are maintenance intensive and necessitate a high weight, since the fluid must be refilled and carried along.
The EMEDS systems have disadvantages in terms of their high weight and with respect to the need for special adaptation of the composite structures to the actuators and their knocking movements. In addition, a de-icing control unit and an energy storage unit are required. Moreover, such systems cannot prevent the formation of ice, but rather can only remove the ice that has accumulated. The removed ice can damage or destroy the engine rotors during the flight.
Exemplary embodiments of the present invention are directed to an ice protection system for an aircraft that avoids at least some of the aforementioned drawbacks.
The invention provides a composite structure for an aircraft component, in order to form a structural element of an aircraft having an ice protection system that is integrated in the composite structure for the purpose of preventing the formation of ice and/or for de-icing. In this case the ice protection system comprises a flat surface-like electric heating element that is embedded in the composite structure below an outer layer of an outer surface of the aircraft component.
Working on this basis, the invention includes a heating unit that is also integrated in a composite structure in order to construct an aircraft component. To this end, a flat surface-like electric heating element is installed below a protective layer, such as a finish top coat, and, as a result, is embedded in the composite structure.
It is preferred that the electric heating element be selected from a group of heating elements that comprise                a textile surface structure of electrically conductive fibers,        a conductive layer made of a carbon fiber-reinforced graphite,        a plastic layer, which is made electrically conductive by filling with graphite,        a conductive layer made of a matrix material that is loaded with nanotubes, and        a layer comprising at least one conductive path that is formed by nanotubes, which are arranged side by side in such a way that they are distributed over the surface of the layer.        
Even more highly preferred is that the heating element is designed as a textile surface structure that can be easily embedded in the composite structure. Preferably the textile surface structure comprises a carbon felt. Such a surface structure can be infiltrated by a matrix material or a cover layer material during application and, as a result, can be almost integrated in the composite structure.
The outer layer is preferably a highly stable thermoplastic material. Especially suitable for this purpose are polyether ketones, for example a polyether ether ketone layer (PEEK). Such materials are easy to shape, on the one hand, and, on the other hand, can be applied in the liquid state and yet are sufficiently temperature stable, in order to embed the heating element in such materials.
The composite structure comprises preferably a carbon fiber-reinforced plastic material as a substrate that forms the support structure for the aircraft component, which is to be de-iced and/or protected against the formation of ice.
Preferably the heating element is connected to the substrate by means of an electrically insulating layer. The layer can be, for example, an insulating layer made of an epoxy material.
When a latent heat storage layer is also incorporated in the composite structure system, the heat of the heating element can be stored in the heat storage layer, so that even after the heating element has been turned off, ice does not form. For this purpose a latent heat storage layer is used preferably with a material that can store heat by means of a phase transition.
Furthermore, the invention provides an aircraft component, in order to form a structural element of an aircraft having an ice protection system integrated in the composite structure to prevent the formation of ice and/or to de ice. In this case the ice protection system comprises a flat surface-like electric heating element that is embedded in the composite structure below an outer layer of an outer surface of the aircraft component.
The electrical contact with the heating element is made by means of at least one contact element that is embedded in the composite structure below an outer layer of an outer surface of the aircraft component. Preferably the contact element is a thermally sprayed contact strip that is electrically connected to the heating element.
Preferably the heating element has a modular type of design and comprises a plurality of flat surface-like heating elements. The electrical contact is made jointly by means of at least one contact element for all of the heating elements or separately for each heating element.
Preferably the method comprises the steps:                providing a substrate made of a fiber-reinforced composite material, in particular a CFC material;        applying an electrically conductive layer, in order to form the heating element;        coating the electrically conductive layer with a cover layer, in order to form the outer surface of the aircraft.        
At least one contact element is applied in order to make electrical contact with the electrically conductive layer. Preferably the contact element is applied after the electrically conductive layer has been applied. Preferably the contact element is coated with a cover layer, in order to form the outer surface of the aircraft.
The electrically conductive layer is applied advantageously in a modular fashion. Preferably at least one contact element is applied between the individual regions of the modular conductive layer.
This is done preferably in such a way that the substrate is coated with an electrically insulating material prior to the application of the electrically conductive layer.
Preferably the electrically conductive layer is formed by or with a material that is selected from a group of materials that include                a textile surface structure made of electrically conductive fibers,        graphite reinforced with carbon fibers,        an electrically conductive plastic due to graphite filling, and        nanotubes, in order to form electrically conductive paths.        